Since the mid 1960s, near-earth space has been populated by ever larger spacecraft, typically today in the ˜1000 kilogram category, launched and boosted into Geosynchronous Earth Orbit (GEO) for communications, or Low Earth Orbit (LEO) for mapping and defense purposes. There is a rapidly growing commercial demand for small satellites of 100-200 Kg mass in sun-synchronous Low Earth Orbit (LEO) of approximately 200 nautical miles, but a significant reduction in the cost of access to orbit for small payloads is essential for success of the emerging commercial space industry. Today's high costs are justified mostly by defense needs, or by launching large satellites into GEO where their useful life is long enough to justify amortized life-cycle costs. Modern satellite launches into LEO use chemical propulsion systems, such as liquid or solid propellant single-stage and multi-stage rockets. Innovative approaches not yet made practical include launching at high altitude from airborne platforms. Advances in structural materials will be one key enabling technology to meeting the challenging cost-to-LEO target. Nano-composite materials could increase the strength/mass ratio of rocket structures and lead to single-stage to orbit with higher payload and hence reduced cost/kilogram delivered to LEO. Exotic chemical propellants that generate higher exhaust velocities would also increase payload delivered to LEO for a given launch pad mass, and hence reduce launch costs.
Small satellites in LEO will place more demands on micro-spacecraft in the 1 kg-10 kg class and small satellites in the ˜100 kg class. With increasing ability to integrate cameras and sophisticated communications systems, the demand for propulsion systems for small (and inherently power limited) spacecraft will grow.
In modern satellites, high thrust for rapid maneuvers has been provided to spacecraft by chemical propulsion, such as hydrazine and other rocket motors. The exhaust velocity of such chemical rockets is limited by the inherent specific energy released by combustion, to ˜2500-3000 m/s. Due to this limited speed, chemical rockets burn up more propellant to effect an orbital maneuver than would other forms of propulsion that offer higher exhaust speeds. These include electro-thermal rockets and electric propulsion. In electro-thermal rockets, the chemical energy released by the propellant is augmented by additional energy input via an external heater. The higher exhaust speeds possible are limited by the temperature at which the rocket nozzle may be safely operated. Electric propulsion is the most efficient in terms of propellant utilization, as it offers much higher exhaust speeds. This is possible because electric rockets add energy to passive propellants via external means and contain the high energy propellant ions or plasma in electromagnetic fields, so that they are not in contact with material walls. Thus the usual limitation on propellant temperature is removed. At high temperatures, exhaust speeds in the 10,000-30,000 m/s range are possible for plasma rockets, while electrostatic ion engines may boost the exhaust speed of ions to still higher velocities, (>100,000 m/s) limited only by breakdown of vacuum gaps at high voltages. Such an order-of-magnitude higher exhaust speed for electric rockets makes them far more efficient in terms of propellant utilization for in-space maneuvers. To illustrate this by example, consider a 100 kg satellite that must be moved in its orbit by a change in orbital velocity of 2000 m/s. If a chemical rocket with 100N of thrust and an exhaust speed of 3000 m/s is used, the orbital maneuver would take about 24 minutes to complete, with a fuel consumption of 49 kg which implies that only half of the initial 100 Kg spacecraft mass would arrive at the destination. By contrast, for an electric propulsion engine with thrust of only 1N, but having an exhaust speed of 30,000 m/s, the same orbital maneuver would take 54 hours but consume only 6.5 kg of propellant, so nearly 94% of the initial mass would be delivered to its destination. The cost/kg of useful payload delivered would be half as much as with the chemical propulsion, in exchange for a longer mission duration. As the required velocity change becomes larger and larger relative to the exhaust speed, chemical propulsion becomes far less efficient. For example, if in the above example, the velocity change were increased from 2 km/s to 4 km/s, the chemical rocket would deliver only 26 kg of the original 100 kg to its destination, vs. 88 kg for the electric rocket. This factor of 3.3 higher useful payload could significantly reduce costs to move objects in space. The above example is illustrative of the general advantage of higher exhaust speed in space. However, the example also shows that the price paid for higher speed electric rockets is often a much longer mission duration, due to the typically much lower thrust offered by such engines, relative to their chemical counterparts. For a given efficiency, the thrust T and exhaust speed are inversely related via:
  T  =            2      ⁢                        P          e                ·        η                    u      exhaust      with Pe being the power into the thruster, η the overall thruster efficiency and uexhaust the exhaust velocity of the rocket engine. As the above example illustrates, chemical rockets have given high thrust but at low exhaust velocity, while electric rockets have given low thrust at high velocities. Orbital maneuvers in space could be dramatically improved if a single propulsion engine were available that offered variable exhaust speed and thrust for a fixed power input at high efficiency. With such an engine, one could operate at high thrust and lower velocity for rapid maneuvers that consume more fuel, but reduce to low thrust at very high velocity, to accomplish slower missions far more efficiently. Rather than carrying two completely different types of engine on board to accomplish this (as is done today) one could utilize a single electric engine to do both tasks.
A new type of electric thruster is known as a Liquid Metal Ion Thruster (LMIT). LMITs offer the advantage that they can be integrated into Micro-Electro-Mechanical-System (MEMS) structures, very similar to current systems being used for field emitters in plasma displays. An LMIT works by producing a high velocity ion current via field emission from a liquid metal source. A high voltage is applied between an extractor electrode at cathode potential and a liquid metal coated field enhancing structure like a small (micron radius) sharp tip. The high voltage leads to the formation of tiny micro tips protruding from the liquid metal surface, known as Taylor cones. These Taylor cones enhance the applied electric field further, leading to a condition where ions can “tunnel” out of the liquid phase into vacuum. The applied extraction voltage accelerates the ions to a velocity u,
      u    exhaust    =                    2        ⁢                                  ⁢        eV                    m        ion            
where:
e as the elementary charge (1.6×10−19 Coulomb),
V is the extraction voltage, and
mion is the mass of the individual ion.
In LMIT systems with increasing extractor voltage, the velocity and the number of ions extracted increase and essentially more thrust is produced.